In the large civil aircraft aviation industry the growth in size of wing tip devices over the years as a result of drive to increase wing efficiency through reduction of drag has lead to technical challenges related to the load transfer and efficient joint technology between the wing tip device and wing. Existing large civil aircraft wing tip attachment methods, such as that described in U.S. Pat. No. 7,975,965 B2, are generally made up of a ‘back-to-back’ rib solution where the loads are transferred across a joint utilising the chord depth of the local wing section.
An innovative solution created to decouple the limitations of load transfer through local chord depth is described in US 2012/0112005 A1. This idea proposes a joint concept that utilises a ‘main beam’ structure to carry the primary wing tip loads and transfer these into the wing via an increased moment arm.
However, the wing tip device tends to be over-engineered, particularly at the attachment point, in order to guarantee the mechanical properties required for the use of such fastening means because current manufacturing methodologies make it difficult to adequately tailor the structural behaviour of the composite beam.
It remains difficult to manufacture and construct using composite materials the complex spar geometry that enables a winglet to be attached to a main wing element. The use of conventional methods such as an assembly of multiple parts to form the spar are difficult due to the lack of access in the geometry available for tooling and assembly, and also inefficient as a result of requiring an increased number of parts, thus increasing cost and weight of the final component, or resulting in a compromise of the structural design to meet the manufacturing constraints.
A known braiding process for forming a complex shaped fibre preform is described in U.S. Pat. No. 8,061,253. The method comprises braiding a plurality of fibres over a non-cylindrical mandrel to form a variable thickness shaped fibre preform. The preform is subsequently flattened and cut to form the spar component. The mandrel is moved at a constant speed during the braiding process.
As noted in J. S. Tate, A. D. Kelkar, and V. A. Kelkar, “Failure analysis of biaxial braided composites under fatigue loading”, The 15th European Conference of Fracture (ECF), Stockholm, Sweden, Aug. 11-13, 2004, when a biaxial braid tube is used for a component of varying cross-section, the braid angle, thickness and areal weight (yield) vary from point to point.
White, Mark L. Development of Manufacturing Technology for Fabrication of a Composite Helicopter Main Rotor Spar by Tubular Braiding. Vol. 1618. KAMAN AEROSPACE CORP BLOOMFIELD Conn., 1981 describes a braided spar for a helicopter main rotor. Each braided layer is designed to be applied at a constant pitch (i.e., mandrel advance per revolution of the braider carriers) allowing the fibre orientation angle to decrease and the layer thickness to increase as circumference decreases along the tapered spar.